Stabilizing surface for aircraft



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STABILIZING SURFACE FOR AIRCRAFT Filed April l, 1947 9 Sheets-Sheet 7 Aug. 7, 1951 w. R. wlNsLow 2,563,298

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satented ug. 7, i-QS UNITED STATES PATENT FFC Y amazes sTABlLlzING SURFACE FoR AIRCRAFT William R. Winslow, North Tonawanda, N. Y. Application April 1, i947, serial No. 738,642

Claims. 1 I. OBJECTS AND GENERAL DEFINITION OF THE INVENTION 1. (a) The invention herein described has for its object an aircraft embodying a new and unique system of stability and control.

This aircraft is composed primarily of a'wing and fuselage combination having a low degree of stability or instability, and attached to and ahead of said wing, an airfoil suspended byl a certain means which causes it to possess a floating action in the airstream.

This floating action is a hitherto known action characteristic of any object, presumably an airfoil, suspended in the airstream by means of a spanwise axis about which there is allowed free rotation. By virtue of this free rotation, the angle of attack of the airfoil remains substantially constant throughout a change in the speed and direction of the airstream.

The term Wing as it is used herein is construed to mean one or more airfoils held in xed or controlled relationship to one another. It is further construed that the wing is a part of the rmain body of the aircraft, to which fuselages, power plants, landing gears, etc. might normally be attached.

When used in combination with the floating airfoil, the wing itself, as defined above, possesses a low degree of stability or instability.

l'n any other known design embodying a floating auxiliary airfoil, the floating airfoil at a given substantially constant angle of attack imparts to the aircraft a pitching tendency of a constant magnitude throughout a change in the main wings angle of attack. As illustrated in Fig. 4, an aircraft having a pitching tendency versus angle of attack as indicated by curve A is effected by this added constant pitching tendency in such a way as to result in a pitching tendency for the complete aircraft as indicated by curve C. It is Furthermore, the pitching tendency at various controlled angles of attack of the auxiliary airfoil imparts a greater stabilizing tendency to the aircraft for higher angles of attack of the auxiliary airfoil. This results in a somewhat greater contribution to static stability at reduced speeds and a smaller contribution to static stability at increased speeds. Referring to Fig. 4, this elfect is illustrated by the greater slope of the air-14 curve for slower equilibrium flight and the smaller slope of the aa=10 curve for faster equilibrium flight.

It is seen to be a necessary requirement of stable flight in other known designs embodying oating auxiliary airfoils for longitudinal control that the aircraft -exclusive of its auxiliary airfoil be possessed of a stable tendency. In the pro.- posed aircraft, a stable tendency of the aircraft exclusive of its auxiliary airfoil, although usually desirable, is not necessary for stability of the complete aircraft. This characteristic further distinguishes this aircraft from other known designs.

The action of the auxiliary or iioating airfoil in conjunction with the unstable wing in free flight is such as to produce inherent stability for the combination of said airfoil and wing.

(b) The invention herein described further embodies a. mechanical means of regulating the aerodynamic force on the auxiliary control airfoil. The free spanwise axis, about which the control airfoil is suspended, has'a movable locato be noted that the slope of curve C is the same y as the slope of curve A. 'I'he static stability is not changed by the addition of the auxiliary airfoil.

The proposed oating airfoil, however, lby virtue of its free pivotal attachment and the locationy of said pivotal attachment in relation `to the aircraft center of gravity imparts to the aircraft a stable pitching tendency which varies inversely with the angle of attack of the main wing. As

indicated by curvev D in Fig. 4, this results in a g bodies a feature by which the auxiliary control more stable v slope of the complete aircrafts {Mam/dd relation. v*The broken line curves in Fig. 4 indicate the dCmoG/da relation for various degrees of longitudinal control, i. e. auxiliary airfoil angles of attack.`

tion relative to the airfoil. This relative location is controllable by the pilot. The designed purpose of this movable relative location is one of enabling the airfoil to be regulated to a variety of angles of attack, such regulation being control# lable by the pilot in flight. Each location of the free axis relative to the airfoil causes the airfoil to assume a certain characteristic angle of attack. This characteristic angle of attack will, by virtue of the floating action, remain substantially constant in night and will change only with a controlled change in the said relative location of the free axis.

(c) The invention herein described has as a further object a means by which the auxiliary airfoils may be manipulated in igh-t by the pilot, thereby to control the attitude of the aircraft in pitch and roll.

(d) A further object of this invention emairfoils may be retracted into the underside of the wing. Iny their retracted positions, the con# trol airfoils contribute to the shape of, and become a part of the airfoil of the wing.

'action of the leading controlairfoils.

(e) A further object of this invention embodies a retracting mechanism which during the extension-retraction motion produces the following conditions relative to the action of the control system: (l) As the control airfoils are near their retracted positions, the aerodynamic forces on the control airfoils are such as to cause the control airfoils to trail in the airstream in a weathervane fashion with small downward (away from the wing) aerodynamic forces. At this position of the control airfoils, the pilots control action at the control stick imparts only a very small degree of control action to the airfoils, and zero control action to the airfoils whenthey are at their fully retracted positions. (2) As the'control airfoils move forward from their retracted positions toward their extended positions, an automatic action which is independent of the pilots control action causes the above described -relative location of the free axis to change in a way as to bring about additional lifting force. This additional lifting force increases in magnitude with extension motion and reaches a desirable positive magnitude at full extension. (3) The effectiveness of the control stick action in producing control action at the control airfoils increases from a zero degree at their retracted positions to a maximum degree at their fully extended positions.

The above involves an auxiliary airfoil designed to travel from the retracted position within the airfoil shape of the wing to an extended position ahead of the wing, and also in a reverse manner back to the retracted position. Accompanying and coordinated with this retracting action there is a further action, referred to in the following paragraph.

(f) A further object of the invention embodies two trailing airfoils which are a part of the main wing of the aircraft and which are located below and to the rear of its trailing edge. In addition to contributing to the lift, the trailing airfoils contribute largely to the instability of the main wing, which must be overcome by the stabilizing The trailing airfoils are designed to be retractable into the airfoil shape of the Wing, thereby to discontinue the main wings unstable tendency. Also, lift and drag are reduced by the retraction o'f the trailing airfoils. This retraction takes place during the retraction of the leading control airfoils. The positive moment produced by the leading control airfoils and the negative moment of the mately to nullify one another both during their coordinated extension-retraction motionl and Yat their fully extended positions. As is characteristic of the mechanical action of the leading control mechanism during extension-retraction, the control stick action in producing action at the trailing control airfoils alsoincreases in effectiveness as the airfoils move toward their extended position. This effectiveness to the motion of the control stick increases from zero at the fully retracted position of the trailing airfoils to a maximum at their fully extended positions. Both the variation in effectiveness to the motion of the control stickand the variation in the lift and moment producing aerodynamic forces (at both leading and trailing control airfoils) are mechanically linked with the extension-retraction motion of the leading control airfoils, 'and are independent of the action of the control stick.

(g) A further object of the invention is a means for vertically trimming the aircraft in flight.

trailing control airfoils are designed approxi- (h) A further object of the invention embodies a device for limiting the variation in the angle of incidence between the leading control airfoil and the longitudinal axis of the main wing. As explained above, this angle of incidence varies in flight as the airfoil floats on the direction of the air in its vicinity. The object of this limiting device is to prevent the airfoil from falling ina pendulum manner due to its weight, and assuming an abnormal flight angle of incidence when the aircraft is travelling at less than its normal ight airspeed.

"(i) A further object of this invention embodies a static balancev (substantially as indicated by part |50 Vin Fig. 2, located along the axis of arm ISI) and. on the side of the free axis of suspension which causes it to counter-balance the Weight of the auxiliary wing, thus serving to nullify the effect of its weight.

(7') A-further object embodiesa device (substantially as indicated by part |52 in Fig. 14) for damping rapid angular motion about the free axis of. suspension.

2. The following is a list of intended advantages claimed in general for an aircraft design embodying lthe invention:

(a) With the auxiliary airfoils in extended positions:

1. Additional lifting surface wing area, of advantage in flying at low speed iiight;

2. Employment of a wingl airfoil shape that is conducive of a high lift coefficient, of advantage in low speed flight;

3. Employment of a fixed slot of highly effective design;

4. Additional control surface area with resulting additional control effectiveness, of advantage in low speed flight;

5. Aerodynamic effectiveness of design wherein all aerodynamic surfaces contribute to the total lift.

6. The feature of positive control action wherein the main `control forces which are located at the leading control airfoils act positively in the direction of the desired control; Y

7 An optional safety feature. depending on an .appropriate aircraft design, whereby a stall on the main lifting surface is prevented by an earlier partiall stall on the auxiliary airfoil.

8. If the invention is used in conjunction with an aircraft designed to hover in a vertical attitude, the arrangement of the control surfaces directly behind the plane of the rotor affords a condition of better aerodynamic control at zero aircraft speed. Also, in this type of design, the elimination of tail surfaces and sweep back in the wing design affords an advantageous condition for the design of'a vertical attitude type landi'ng gear (such as illustrated in Fig. 20).

9. Asafety feature embodying a stall warning manifested by a characteristic oscillating motion of Vthe auxiliary airfoil about its` axis of suspension as it approaches a stall.

(b) With auxiliary airfoils retracted:

Low 'Wing area, desirable for high speed flight;

2. Employment of a nearly symmetrical wing yairfoil shape conducive of low `drag and of an In the drawings accompanying and forming part of this disclosure- Y Fig. l is a diagram of the main airfoil and auxiliary airfoil showing the areodynamic force coecients in their relative locations;

Fig. 2 is a diagram of the auxiliary airfoil and supporting structure showing the resultant force vector Cna for (aa=l2) and broken line vectors for (aa=8) and (aa=16).

Fig. 3 is a table of airfoil data and assumed proportions of design used in the stability calculations.

Fig. 4 is a plot of the aircraft pitching moment CmC G versus the wing angle of attack (aw) for (A) the main wing alone and (D) for the combined main wing and auxiliary wing.

' Fig. 5 and Fig. 5a, respectively, show the eiect of low and high velocity on angle of attack of an airfoil in free flight.

Fig. 6 and Fig. 6a, respectively, show the effect of low and high velocity on angle of attack of an airfoil suspended in flight by a spanwise axis about which there is allowed free rotation.

Fig. 7 is a diagram of the combination of an unstable wing and a freely suspended auxiliary airfoil, showing the effect on the pitching tendency brought about by an airspeed which is lower than the equilibrium air speed for the given angle of .attack of the auxiliary airfoil.

Fig. 8 is a diagram of the combination of an unstable wing and a freely suspended auxiliary airfoil, showing the eiect of the pitching tendency brought about by an airspeed which is higher than the equilibrium air speed for the given angle of attack of the auxiliary airfoil.

Fig. 9 is a diagram of the control system in the extended configuration of the retractable control surfaces.

Fig. 9a, is a diagram of the control system in the retracted configuration of the control surfaces.

Fig. 9b is a diagram of the relative airfoil shapes of the auxiliary airfoil and the airfoils comprising the wing in the extended conguration.

Fig. 9c is a diagram of the relative shapes of the auxiliary airfoil and the airfoils comprising the wing in the retracted configuration.

Figs. 9d, 9e and 9j are diagrams of the relative positions of the primary mechanical parts which support the trailing airfoil of the wing showing, respectively, three positions: (d) fully retracted, (e) intermediate and (f) the neutral control position corresponding to the fully extended position of the auxiliary airfoil retracting mechanism.

Fig. 10 is a diagram showing the primary mechanical parts of the control system in the extended configuration.

Fig. 10a is a cross sectional diagram of beam 4| and rollers 42 and 43.

Fig. 11 is a diagram showing the primary mechanical parts of the control system in the retracted configuration.

Fig. 12 is a schematic diagram showing typical electrical wiring land mechanical parts for the transmission of motion between rods 2, |22, I 2|, 22, and their respective control surfaces.

Figs. 12a. and b are auxiliary views of cam 4 and segment 5.

Fig. 13 is a diagram of the retracting mechanism of the auxiliary airfoils 54 and 55 in the extended configuration.

Fig. 13a is a diagram of the retracting mecha-v between the retracted and extended conlgurations of the auxiliary airfoil.

Fig. 13d is a diagram showing the relative positions of the primary vmechanical parts of the retracting mechanism in the extended configuration of the auxiliary airfoil.

Fig. 14 is a diagram showing the primary mechanical parts of the retracting mechanism in the extended conguration.

Fig. 14a. is a detail cross section sketch of carriage 8'ltrack 9i and rollers 89 and` 9U at axis I. v l

Fig. 15 is a diagram showing the primary mechanical parts of the retracting mechanism in the retracted configuration.

Fig. 16 is a schematic diagram showing the electrical wiring and mechanical parts between the retracting control switch 59 and motor 1U.

Fig. 17 is a schematic diagram rshowing the electrical wiring and mechanical parts for the transmission of motion between motor 10 and motor |05.

Fig. 18 is an isometric view of the cockpit control unit.

Fig. 19 is a detail View of the mechanism for coordinating the control action with theV retracting action showing a cross-section in the plane of axesP and Q.

Fig. 20 is an auxiliary side view along the line 2li- 20 of Fig. 19 showing a mechanism at axes P, Q, R and S.

Fig. 21 is a .side view as viewed from the left of the cockpit control unit with parts of frame I removed to show the working parts.

The invention will be best understood by first considering the aerodynamic theoryV upon which it is based, with particular respect to an analysis of stability.

1I. ANALYSIS OF STABILITY l. Quantitative analysis of stability assuming constant 'velocity and zero acceleration Two airfoils of known aerodynamic characteristics are used in the following analysis. The two airfoils used are a U. S. A. 35B airfoil for the main wing and an N. A. C. A. 0012 airfoil for the auxiliary wing.

Abbreviations applying to Figs. 1 to 8:

CMM- moment coefficient about the aerodynamic center. rif-approximate distance in percent of M.'A. C.

p-air density.

lift of the main wingf` t IgE-litt o .fgthe auxiliary; airfoil. wnweight., o fg the aircraft. iv-ansie of., attack. main wins. a--angle of attack, auxiliarywing; SVA-planiform area ofthe mainfwing. Saz-1:21aniformV area. of theauxiliaryl Wing.`

Referring-to Fig: 1, aV summation of moments istaken about the aircraft center of' gravity;

An, angleof-i attack isassumed for the auxiliary airfoil which, by virtuev of its free span-wise axis of; suspension, remains 4constant under conditions of.' varyingwing.` angle: of attack;

Two conditions of wing angle ofv attackare assumed, one higher and one lower than the angle of' attack for equilibrium ilight.

Referring to known airfoil characteristics, the values of aw, CLw, C. P, an, CRa and approximate df/M; A. C; arercompiled in Fig. 3.

Aircraft design proportion Sii/Sw is assumed toLbe11/4. Thelcenter of gravity -is assumed to be located at- .25 M. A. C.

SolvingEquation l1 for CmC G at the condition ofithe: higher aw, Y

Solving-.Equation 11 for Gmc@ at .the condition of the lower aw,

` C'mQ.G =-.08.69(.36-.2 5 +(.90). (.8) (1/4).

CmCG` =..024"

The slope of the CmC G curve,l thus obtained and" illustrated bycurve D.: in Fig. 4;.indicates1a stable.. tendency.

2. Qualitative analysis of static stability assuming freeflight conditions The foregoing explanation andv determination of stability isgfoundedon agbasis ofydisplacement oithefwing .angle ,of attackV withfairspeedfassumed to .remain constant.

Under actual conditions of free flightiinisrnoothi air, airspeed and altitude are added variables. A normally stable airplane disturbed from a. condition of` equilibrium'. will oscillatorY inV av verticalA plane..withairspeed-varying inversely with altitude, and with oscillations diminishing .in magnitude until' aY condition of equilibrium is again.

reached; During vsuch stable oscillations the angle ofattackof-the main wing varies inversely with thesquare of` the velocity.

a4 constant- OL :T1-TF* CL-x Therefore The angle f of attack; of the; auxiliary.; airfoil, freely suspended ahead of the wing, however, remains substantially:the.same/throughout:avaria: tion ,in the speed ,and direction ,ofthev airstream.

Infthefollowingl explanation it; will be., seen-1 that changes in airspeed contribute a largedegvree Itoward thefstabilitm An increase inthe airspeed is.,,accomDanied by a decrease in the wing angle of attackrwhilegthe..

liftzon; thawing; remains.. substantially. the.; same. (See Figs. 5 and 5a.)

Atethpe; sameft-imaoomY the auxiliary airfoila this increaseI inairspeedfA brings about.l a` pronounced increase indift: (Figs. 6'Yand6ag) dnces aclimbngtendency and. thereby ay loss of airspeed;-

A` lfurther v effect which contributes toward stability, notlmentioned` heretofore, is the changingv value of which canbe seen below todecrease with .an increase in La.

2F11.: 0, W=constant :Liri-Lwv Lav-:constant-L.

An increased Lw inFig. ''further contributesl toward; a1 diving tendency.Y A decreased Lw in Fig- 8" furtherl contributes toward a climbing tendency.'

3.. Analysis oj'vstability injree .flight under a gust.v

` condition A\further1free' flightcondition involving an Voutsideldisturbance -such-as-a' vertical gust mayfin duce a change in the angle of attacky without changingzthemirspeed,

It: isasupposedlthat the aircraft'. encounters an' outside disturbance such as a gust which suddenly. increases the wing angle-of attack-'without decreasingcthe' speed;` With'this increase in `wing anglelof: attack; the--liftaon'the wing immediatelyA i increases.

Since the new lift force is in excess: ofvtlie` value necessary to support-the weightof the'. aircraft;.. the aircraft. accelerates upward; Two stablevariations .are hereby. brought aboutzr (l) An*L increased. lift; force. located behind the center ofgravity. and an added downward accelerationxforce atv the center. ofrgravity constitute' azforcecouple which produces a diving. moment.

(2.);An1unward vertical Velocity. brought. about;v

by.f thay upward` acceleration. of the aircraft.'

changesT the-1 direction.. of. the air. striking; theT wing in such a way as to decrease the angle of. attacknf thawing.

Oppositely,,stableltendencies may be. seen, to.. be brought about by a sudden1decrease in angle.

\- of attack without an` increase in airspeed.`

Therefore, the immediate effects of;asudden.

change in angleof attack without change in airspeedj areV a.. stable. pitching tendency and. a` change in angle of attack in thedirection ofthe.

t original angle of attack- IIL. EiX1E.I1A1\IATIONv OFA MECHANICAL EUNCTIONS l.v Control mechanism The control mechanism is substantially: asn

shownginFigs-9, 10,11; l2 and-18..

. Referring;` to Fig,j 10f,thecontrol. stick, I is. moved in the directions indicatedzby,theuarrows..

75 pointingv up; down. tocthe leftandltmthezrizht.

These motions impartthe same corresponding control motions to the attitude of the aircraft in flight. For example, an upward motion of the control stick produces a nose-up change in the attitude of the aircraft. A motion of the control stick to the right produces an approximately coordinated bank and turn to the right.

In the detailed explanation which follows, certain mechanical changes as described are caused to be brought about by certain motions of the control stick.

For the sake of simpueity, only the left-handv chanical changes in the sequence given are brought about by an upward motion of the control stick. The control stick and frame I rotates about axis A, moving rod 2 forward and rotating arm 3 about axis C. Cam 4 is mounted on an axis which lies parallel with the major axis of arm 3. About this axis, the cam is free to rotate against the force of spring I54 as illustrated in Fig. 12b which tends'to return it to its angu-v lar position, as shown in Fig. 12a. kAs arm 3 rotates about axis C, electrical terminals 8 and 9 on cam 4 connect with resistances I2 and I3, respectively, thereby allowing electrical current to pass from the resistances through the terminals to ,the D. C. motor I4. Motor I4 which is stationary with the structure of the auxiliary airfoil and the worm gear I5 rotate gearsegment I6 and impart a counter-clockwise rotation (as viewed from the left) to shaft I1, causing a change in the relative location of the axis E with respect to the auxiliary airfoil. The rotation of shaft I1 is transmitted through gears I8 and I9 to Selsyn 2U. Selsyn 2! is connected to segment 5. Selsyns 20 and 2I are electrically interconnected so that the angular position of shaft I1 is transmitted by Selsyns 20 and 2I to segment 5. Segment 5 rotates until the electrical contact between terminals 8 and 9 and resistances I2 and I3 has been broken, thereby causing D. C. motor I4 to stop.

In a similar manner, an opposite or downward motion of the control stick causes electrical contact to be made between terminals 5 and 1 and resistances I and II, thereby transmitting electrical current of a reverse polarity to motor I4 and causing the motor to move axis E in the opposite direction. Also, as described above, the Selsyns cause segment to follow and overtake the position l of arm 3.

Ina manner similar to that of the above description; the same upward motion of the control stick also causes rod 22 to rotate arm 24, motor 34, worm gear 35, gear 3S and shaft 31.v Selsyns 38 and 39 complete the action as described above. Shaft 31 imparts a clockwise (as viewed from the left) rotation to arm 4l). Axis G moves downward and` forward, carrying with it beam 4 I. Beam 4I is suspended at axis G by two rollers 42 and 43. The two rollers are conned to two channels, one on each side of beam 4I. As axis G moves downward and forward, beam 4I rotates and rolls forward'perpendicularly through axis H, imparting an upward and forward motion to the trailing edge of control surface 44. -It may be seen also that a downward motion of the Conf trol stick brings about a rearward and downward motion of the trailing edge of control surface 44.

The upward motion of the control stick as described above also moves rod 45 forward. This motion is transmitted through bell crank 46, rod

41, bell crank 48, rod 49, bell crank 5B, rod 5I tocontrol surface 52, causing its trailing edge to move upward. In a similar manner, a downward4 y motion of the control stick causes to be brought about a downward motion at the trailing edgev of control surface 52.

A vector diagram in Fig. 2 shows resultant` aerodynamic forces and their relation to the airfoil of the control surface. The length of arm I5I between axes D and E together with the loca- 1 tion of axis D in relation to the airfoil are determined such that the free axis E when rotated about axis D will intersect the CRa vectors at in-,

tervals along its path of motion which areas evenly spaced as possible. These points of intersection are designated on Fig. 2 by the angles ofV For exam-y ple, the point of intersection of the axis E and theattack corresponding to the vectors.

vector CR,L at aa=12 is designated 12.

Suppose that the axis E whose location rela-l tive to the airfoil may be changed by the pilot be given the location shown in Fig. 2. At this location, the resultant aerodynamic force on the airfoilpasses through E. Consideringthis force as the only force acting on the airfoil which is freel to rotate about E, the airfoil is seen to be in equilibrium. The angle between its chord line and the direction of air through which it is passing is 12. The airfoilisfree to rotate about'axis E and its angular relationship about axis E is dependent only on the airdirection.

If the air direction should change, it may be seen from the vector diagram that the airfoil will rotate about axis E tocause its angle of attackv Now suppose that an.

of 12 as described above. upward gust strikes the airfoil and causes its angle of attack to increase momentarily to 16. At a 16 angle of attack the resultant aerodynamic force is represented in Fig. 2 by Cna at aa=16. Itmay be seen that this force vector passes to the rear of the axis E. This force acting on the airfoil times its perpendicular distance to If the angle between arm I5I and the airfoil.

were increased runtil axis E coincided with the vector Cea at 11:16, the airfoil would be caused to float at 16 as described above.

From the vector analysis of the floating action shown in Fig. 2, it may be seen that for an upward motion ofthe control stick, and a resulting counterclockwise rotation of the free axis E about axis D, the two vcontrol surfaces 54 and 55 located ahead of the center of gravity of the aircraft are caused to float at an increased angle of attack and acquire additional lift force, while the four control surfaces located behind the center of gravity 44, 45, 52 and 53 move toward positions which reduce their lifting force. The combined effect of these changes in lift brought about by an upward movement of the control stick is such as to rotate the aircraft about its lateral axis and cause it to assume an angle of ammo 11v steeper climb. Oppositely, 'a downward `motion of the control stick may Hbc seen to -cause the aircraft to dive.

To control the aircraft inthe Ainitiation of a normal turn to the right, ythe control stick :is moved to the right as indicated by the arrows in Fig. 10. rlhe control stick and frame rotate about axis B, pushing rod '2 forwardand pulling rods 22 and 45 to the rear. The three motions of rods 2, 22 and-45 -produce the following three respective motions -at-axis-E, control surface 44 and control surface 52. Axis E moves downward with respect to the airfoilshape of the leading-control surface 54; the trailing edge of control surface 44 moves to -the `rea-r and downward; the trailing edge of control surface 52 moves downward. It may -be seen here that `all control surfaces 44, 52 and 54 tothe left ofthe aircraft center line change positions in manners-such as to produce additional lift. It may also be seen from the mechanism shown in Fig. 9 that all control surfaces 45, 53 and 55 to the right of the aircraft centerline change vpositions in mannerssuch as to reduce lift. The combined effect of these changes in lift y'brought about Yby a movement of the control stick to the right causes the aircraft to roll to the right about 'its longitudinal axis.

A further effect of the control stick vmotion to the right is brought vabout by the negative dihedral angle of control surfaces 52 and 53. (See Fig. k9.) This 'negative dihedral angle which heretofore has been applied effectively .to other designs causes the control surface ,action of a roll 4to the right kalso to produce a yaw to the right. The proportional amount of yaw to roll is governed by the :magnitude of the negative dihedral angle. The negative dihedral angle of the outer portions of the wing is therefore designed to produce the combined proportion of yaw and roll desirable in a normal turn.

It is also to be noted here that a positive dihedral angle of the leading control surfaces 54 and 55 serves to produce a similar tendency to yaw'in the direction of roll. Here also the magnitude of the designed dihedral angle may be such as to Vproduce more or less yawing tendency with roll.

2. Leading control surface free afngular displacement limiting device The angular displacement of a leading .control surface 54 or 55 about its free spanwise axis of suspension varies with the changing direction of the surrounding airstream. 'In normal flight at any one control position of vthe free axis with respect to the control surface, the control surface varies throughout a certain normal range of angle of incidence tothe structure which attaches the free axis to the main body of the aircraft. At another such control position there is another such normal range of angle of incidence. It is the purpose of the limiting device to limit, for each .control position, the normal range of the said angle of incidence to such a range that is necessary for and within which the angular position of the control surface might occur in normal flight.

Referring to Figures 2-and 14, the limiting device consists of a cylinder |52 which is attached to either one of the two axes M and N. To the other one of these two axes is attached a connecting rod and piston |53 which is free to move longitudinally through cylinder |52. 'By limitation of the relative motion of the piston and cylinder, it Vmaybe seen from the figure-that the rotation yof the control surface about axis E is also limited. Therefore a desirable 'limitation of the angular motion about the Afree axis E for 2a certain control position of arm |5| is accomplished by a certain limitation of the relative linear motion between the piston |53 and cylinder `152.

By properly designing the distances between axes E and M and between D and N, the limited range of the angle of incidence is caused to change -with a change in the controlled position of arm |5I. This change is such that for each control position, the normal night range of the angle of' incidence -will be within the range limited by the device.

3. ,Leading .control surface free .angular motion `dmziping device This Adeviceis a means for damping .or reducing excessively Arapid angular motion .about .the free axis of .suspension of the leading control surfaces. The device lconsists primarily of piston |53 and cylinder |52. It is provided that a viscous fluid .be contained in the cylinder head 4and that port .holes in thepiston be provided for Ythe passage of the viscous fluid through the piston.

.By this means, the internal friction-of the viscous fluid is used to reduce-excessively rapid motion 'of the piston in the cylinder and, by the mechanism explained in 4(2) above, therefore to reduce Yexcessively rapid angular motion about the free axis E.

A. .Control .surface retracting mechanism The .control surface retracting mechanism is substantially as shown in Figs. 13, 14, 15, 16, 17 and 18. The position of control surfaces ,44, 45, 54 and 55lbetween the two positions shown in Figs. 13 and 13a .is .controlled by .the pilot by means of .electrical toggle switch 59,. (See Fig. 14.) The action of the mechanism in moving the control surfaces from their extended position to their retracted positionis as follows:

.Referring to Fig. 16, the toggle switch 59 is turned to 4the position indicated Retract Direct current flows in .conductors 60 and 6| to conductors .Bland B5 on the forward face of insulated .disc 63, .then through conductors 61 and B8 to D. C. motor 1U. Motor 10 turns worm gear and disc 63. Motor l0 valso turns gears '|2 and 13, shaft '|4 `and the following parts which are typical lof the four retracting jacks: bevel gears 1.5, 16, 'l1 and '18,.shafts 19 and 8|), gears 8|, 82,83, 84 and retracting screws .85 and 86.

The rotation of the retracting screws moves carriage 81 to the rear. Carriage 8l carries with it beam 88 and rollers .89 and 90 which rotate about axis I. The transverse motion of rollers 89 and 9U is confined to two channels in track 9| Connected to ybeam 88 at axis J is strut 92. As axis I moves to the rear along .track 9|, axis J and strut .92 rotate about axis K. The leading control surface moves .downward and to the rear, and then upwardinto its retracted position within the airfoil shape of the Wing. The lengths of the circular conductor strips on disc 63 are designed in a way such that electrical contact between conductors 6| and 65 will be broken as the control surfaces reach their retracted position. Unthreaded surfaces on screws and 86 (similar in construction 4to screws H3, I4, ||5 and IIS shown in..detall in Fig. 18e) `and springs similar to "springs 12| and |22 are designed to prevent carriage 81 lfrom overriding its limiting positions-' both at full extension and at full retraction.

To move the leading rcontrol surfaces from their retracted positions to their extended positions, the toggle switch 59 is turned to the position marked Extend D. C. current flows through conductors 6| and 62, 65 and 66, and 68 and 69,

causing D. C. motor 10 and mechanical parts 13Y t'o 92'to operate in a reverse direction and causing leading control surfaces 54 and 55 to move to their fully extended positions. As they reach their position of full extension, electrical contact between conductors 6| and 62 and conductors 65 and 66 is broken to cause D. C. motor 10 to stop.

It is to be noted here (1) that the retracting action may be stopped and started in any position or reversed in either direction between full extension and f-ull retraction; (2) that the extension-retraction action stops automatically when leading control surfaces 54 and 55 reach their fully retracted or fully extended position; and (3) that at -both the limiting extended and at the limiting retracted positions of leading control surfac-es 54 and 55, the toggle switch is effective only in causing motion which is in a direction away from these limiting positions.

Disc 63 has two faces, each of which are shown respectively in Figs. 16 and 17. Referring to the rear face in Fig. 17, the above described motion of the disc during retraction causes disc 63 and cam to come in contact with one another and to make an electrical connection respectively between terminals |0| and |02 and resistances |05 and |06, allowing current to flow to D. C. motor |05. D. C. motor |05 turns shaft |06 (Fig. 18)

and gears |01, |08, |09, ||0, and ||2. Screw.

shafts ||3, H4, ||5 and ||6 rotate in unison and carry respective keyed fittings |.|1,.| I8, ||9 and toward the inner extremities of frame The mechanism is designed so that fittings ||1 to |20 will reach their respective ends of threaded portions of shafts |3- to I6 at approximately the same time that the leading control surfaces reach their retracted positions. The unthreaded portions of shafts i3 to |6 are designed to prevent fittings from overriding their limiting positions. Springs |2| and |22 which are typical of springs located at both ends of shafts ||3 to ||6, serve a` designed purpose of starting the fittings ||1 to |20 onto the threaded portions of shafts ||3 to I6 when their rotationy is in a direction to cause the fittings to move away from their limiting positions.

Fig. 18 and Fig. 14 show the fittings ||1 to |20 in positions corresponding to the extended positions of leading control surfaces 54 and 55. Fig. 15 shows the mechanism with these fittings in positions corresponding to retracted positions of leading control surfaces 54 and 55. In the latter positions, rods 2, |22, |2| and 22 pass approximately through the intersection of axes A and B. Therefore all control stick motion is ineiective in producing motion at any of the four retracted control surfaces 44, 45, 54 and 55. With this con,

figuration of the control system, control of the aircraft is maintained by means ofthe outboard trailing control surfaces 52 and 53 whose allied mechanisms remain unchanged by the retracting action.

' Referring to Fig. 1'1, motor |05, in addition to providing rotation of shaft |06, rotates Selsyn |26 through a reduction gearing |24. Connected to the axis of Selsyn |21 is arm |28 upon which cam |00 is suspended free to rotate about an axis which lies parallel with the major axis of arm |20.

current to motor |05. Therefore, with the mech-- anism as shown, the retraction motion of the control surfaces is synchronized with the motions.' of fittings ||1 to |20 with respect to the control frame The relative sizes and positions of the allied l parts ofthe retracting and control mechanisms. particularly the positions of Selsyns 2|, |30, |3|`` and 39 with respect to the control frame are designed so that the following conditions will hold true during the extension-retraction mo` tion of the control surfaces: (l) Control eifective" ness will increase from zero at the configuration' shown in Fig. 13a to a maximum degree at the- (2) As the auxil-i configuration shown in Fig. 13. iary control surfaces 54 and 55 are at or near their retracted positions, the position of axis E relative to the airfoil of either control surface issuch that the control surface will trail in the airstream with a small downward. (away from the' wing) aerodynamic force. At the same time, thel trailing control surfaces 44 and 45 will be in posi-- tions ycausing them also to produce downward It is an object of the design of the control system to cause the leadingl and trailing yairfoils to -be inclined in the air-S aerodynamic forces.

stream at angles such that their pitching moments produced by their Ydownward aerodynamic forces will approximately nullify one another. As

the leading and trailing control surfaces move further. in the direction of extension, these twodownward forces diminish and are replaced by increasing upward aerodynamic forces. At the fully extended positions of the leading and trailing control surfaces, their respective moment'effects inpitch are designed to a desirable degree to nullify one another. A discrepancy in the balance of those moments would give the aircraft a This tendency mayf be overcome either by an intentional movement J t of the control stick or by a manipulation. of a'- trimming device which is explained in the fol-u tendency to climb or dive.

lowing paragraph.

5. Aircraft vertical attitude trimming device rIlhe trimming device herein described is designed to enable the pilot to trim the vertical at titude of the aircraft and thereby to reduce persistent control pressures and to preserve a maximum of up and down control movement from*v This device is designed to be effective in regulating thecontrol position at horizontal flight.

trim at either the extended or the retracted con--v iigurations of the control system. Y

The adjustment of trim Yis accomplished manually by means of hand crank |40. (See Fig. 2l.) The hand crank is pulled forward and sideward to release it from retaining hook I4 Spring |42 rotates structure |43 about axis L and forces bevel gear |44 and ,idler gear |45 to the rear, com-` pressing spring |46. This action slides bevel 'gear |01 to the rear, as indicated in Fig. 19 byarrows, along its keyed shaft, disengaging it from gears |08 and |09, and engaging gears |45 with gears |08 and |09. A rotation of the hand crank prothe vdesired trim adjustment has been made, the

hand crank is pulled -forward and engaged with retaining Yhook IM. This allows spring V|46 to re-engage bevel gear i`| with gears i08 and |09,

such gear `positions being ,necessary for the n'ormal 4operation of the extension-retraction mecha- What is claimed is:

L1. An aerodynamic control device automatically maintaining an angle of attack irrespective of changes `in air direction, comprising in combination with a main wing and laterally 'spaced struts projecting forwardly therefrom, larms `pivotally supported for relatively free movement about a lateral axis fon the forward ends of said struts, van auxiliary airfoilpivotally Asupported on a ilateral axis on the Afree ends of said pivotally supported arms, saidlatter axis 'extending spanwise through the central 'portion of the airfoil and remote controllmeans for varying the angular relation `between said supporting arms and the airfoil pivotally supported thereby and -for maintaining the same in selected angular relations irrespective of -said -relatively .free pivotal movement of the supporting arms on the forward ends'of saidstrut's.

An aerodynamic ycontrol vdevice automatically maintaining an angle of attack irrespective of `changes in air direction, comprising in combination with va rmain wing and laterally spaced struts projecting forwardly therefrom, arms pivotally supported for relatively free `movement about a lateral axis onthe forward ends of said struts, lan. auxiliary airfoil pivotally supported on a lateral 'axis von the free ends of said pivotally supported rar-ms, said latter axis extending spanwise through the central lportion 4of the airfoil and remote controlmeans for varying the angular relation `between said supporting arms and `the airfoil pivotally supported thereby and for maintaining the 'same in selected angular y relations irrespective of said relatively free pivotal movement of the supporting arms on the forward ends of said struts, and damping means carried by at least one of said struts nfor controlling the lrela-- tive .freedom of motion of said .pivotally `supported arms and Aairfoil carried thereby.

73. An aerodynamic control device automaticallywmaintaining anangle of attack irrespective oi changes in air direction, comprising in combination with a-main wing and laterally spaced struts projecting forwardly therefrom, arms pivotally supported for -relatively free movement abouta lateral axis on theforward ends of said struts, an auxiliary airfoil pivotally supported on a lateral `axis on the free ends of said pivotally supported arms, saidlatter axis extending spanwise through the central portion of the airfoil and remote control means for varying the angular relation between said supporting arms and the airfoil pivotally supported thereby and for 16 maintaining the same in selected angular relations irrespective of said relatively free pivotal movement of the supporting arms on the `forward ends of said struts and means for effecting the forward projection and rearward retraction of said struts with respect to said main wing.

4. An aerodynamic control device automatically maintaining an angle of attack irrespective of changes in air direction, comprising in combination with a .main wing and laterally spaced struts projecting forwardly therefrom, arms pivotally supported for relatively free movement about a lateral axis on the forward ends of said struts, an auxiliary airfoil pivotally supported on a lateral axis on the free ends of said pivotally supported arms, said .latter axis extending spanwise through the central portion of the airfoil and remote control means for varying the angular relation between said supporting arms and the airfoil pivotally supported thereby and for maintaining the same in selected angular relations irrespective of saidrelatively free pivotal movement of the supporting arms on the forward ends of said struts, means for effecting the vforward projection and rearward retraction of said struts with respect to said main wing and means for automatically changing the angle of said airfoil with respect to said arms in accordance with the extent of projection and retraction of said struts.

5. An aerodynamic control device automatically maintaining an angle of attack irrespective of changes in air direction, comprising in combination with a main wing and laterally spaced struts Aprojecting forwardly therefrom, arms pivotally supported for relatively free movement about a lateral axis on the forward ends of said struts, an auxiliary airfoil pivotally supported on al lateral axis on the free ends of said piv- I otally supported arms, said latter axis extending spanwise through the central portion of the 'airfoil and remote control means for varying the angular relation between said supporting arms and the airfoil pivotally supported thereby and for maintaining the same in selected angular relations irrespective of said relatively free pivotal movement of the supporting arms on the forward ends of said struts, means for effecting the forward projection and rearward retraction of said struts with respect to said main wing and means for automatically changing the angle of said airfoil with respect to said arms in accordance with the extent of projection and retraction of said struts and including leverage connections for rendering said remote control means completely effective when said struts 'are fully projected and less effective as said struts are retracted.

WILLIAM R. WINSLOW.

REFERENCES CITED The following references are of record in the file of this patent:

UNITED STATES PATENTS Number Name .Date

l1,806,379 Wood May 19, 1931 1,837,132 Page -Y Dec. 15, 1931 1,862,902 McDonnell e; June 14, 1932 2,049,188 Alfaro L July 28, 1936 2,156,994 Lachmann May 2, 1939 2,198,893 Van Vaveren Apr. 30, 1940 v2,428,194 Bockrath Sept. 30, 1947 

